Heliplane rotor thermal management for maintaining dimensional stability

ABSTRACT

A rotor system is disclosed for a reactive drive rotary wing aircraft. Apparatus and methods are disclosed for maintaining the rigidity of the rotor and eliminating play between flight controls and the rotor by mounting swashplate actuators to a flange rigidly secured to the mast. Apparatus and methods are disclosed for thermal management of the rotor in order to avoid bearing failure or loss of bearing preload. Methods include modulating the temperature of oil pumped over one or more of the mast bearing, swashplate bearing, and spindle bearing. The temperature of air passively or actively drawn through rotor may also be modulated to maintain bearing temperature within a predetermined range. Structures for reducing pressure losses and drag on components due to air flow through the rotor are also disclosed. A rotor facilitating thermal management by oil and air flow is also disclosed.

RELATED APPLICATIONS

This application: claims the benefit of U.S. Provisional PatentApplication Ser. No. 61/403,111, filed on Sep. 9, 2010. This applicationincorporates herein by reference U.S. Provisional Patent ApplicationSer. No. 61/381,291, filed on Sep. 9, 2010, U.S. Provisional PatentApplication Ser. No. 61/403,099, filed on Sep. 9, 2010, U.S. ProvisionalPatent Application Ser. No. 61/403,097, filed on Sep. 9, 2010, U.S.Provisional Patent Application Ser. No. 61/381,313, filed on Sep. 9,2010, U.S. Provisional Patent Application Ser. No. 61/403,111, filed onSep. 9, 2010, U.S. Provisional Patent Application Ser. No. 61/381,347,filed on Sep. 9, 2010, U.S. Provisional Patent Application Ser. No.61/403,136, filed on Sep. 9, 2010, U.S. Provisional Patent ApplicationSer. No. 61/403,134, filed on Sep. 9, 2010, U.S. Provisional PatentApplication Ser. No. 61/460,572, filed on Jan. 3, 2011, U.S. ProvisionalPatent Application Ser. No. 61/403,098, filed on Sep. 9, 2010, U.S.Provisional Patent Application Ser. No. 61/403,081, filed on Sep. 9,2010, U.S. Provisional Patent Application Ser. No. 61/403,135, filed onSep. 9, 2010, U.S. Provisional Patent Application Ser. No. 61/466,177,filed on Mar. 22, 2011, U.S. Provisional Patent Application Ser. No.61/409,475, filed on Nov. 2, 2010, U.S. Provisional Patent ApplicationSer. No. 61/403,113, filed on Sep. 9, 2010, U.S. Provisional PatentApplication Ser. No. 61/409,478, filed on Nov. 2, 2010, U.S. ProvisionalPatent Application Ser. No. 61/409,476, filed on Nov. 2, 2010, U.S.Provisional Patent Application Ser. No. 61/409,482, filed on Nov. 2,2010, U.S. Provisional Patent Application Ser. No. 61/409,470, filed onNov. 2, 2010, U.S. Provisional Patent Application Ser. No. 61/517,413,filed on Apr. 19, 2011, U.S. Provisional Patent Application Ser. No.61/468,964, filed on Mar. 29, 2011, U.S. Provisional Patent ApplicationSer. No. 61/409,487, filed on Nov. 2, 2010, U.S. Provisional PatentApplication Ser. No. 61/409,494, filed on Nov. 2, 2010, U.S. ProvisionalPatent Application Ser. No. 61/456,219, filed on Nov. 2, 2010, U.S.Provisional Patent Application Ser. No. 61/456,221, filed on Nov. 2,2010, U.S. Provisional Patent Application Ser. No. 61/456,220, filed onNov. 2, 2010, U.S. Provisional Patent Application Ser. No. 61/432,488,filed on Jan. 13, 2011, U.S. Provisional Patent Application Ser. No.61/506,572, filed on Jul. 11, 2011, U.S. Provisional Patent ApplicationSer. No. 61/519,075, filed on May 16, 2011, U.S. Provisional PatentApplication Ser. No. 61/519,055, filed on May 16, 2011, U.S. ProvisionalPatent Application Ser. No. 61/460,573, filed on Jan. 4, 2011, U.S.Provisional Patent Application Ser. No. 61/461,223, filed on Jan. 13,2011, U.S. Provisional Patent Application Ser. No. 61/429,282, filed onJan. 3, 2011, U.S. Provisional Patent Application Ser. No. 61/429,289,filed on Jan. 3, 2011, U.S. Provisional Patent Application Ser. No.61/499,996, filed on Jun. 22, 2011, U.S. Provisional Patent ApplicationSer. No. 61/575,196, filed on Aug. 17, 2011, and U.S. Provisional PatentApplication Ser. No. 61/575,204, filed on Aug. 17, 2011, all of whichare hereby incorporated by reference in their entireties.

Additionally, this patent application hereby incorporates by referenceU.S. Pat. No. 5,301,900 issued Apr. 12, 1994 to Groen et al., U.S. Pat.No. 1,947,901 issued Feb. 20, 1934 to J. De la Cierva, and U.S. Pat. No.2,352,342 issued Jun. 27, 1944 to H.F. Pitcairn.

RIGHTS OF U.S. GOVERNMENT

This invention was made with Government support under Agreement No.HR0011-06-9-0002 awarded by DARPA. The Government has certain rights inthe invention.

BACKGROUND

1. The Field of the Invention

This invention relates to rotating wing aircraft, also known asrotorcraft, and, more particularly to rotating wing aircraft relying onautorotation of a rotor to provide lift.

2. The Background Art

Rotorcraft rely on a rotating wing to provide lift. In contrast, fixedwing aircraft rely on air flow over a fixed wing to provide lift. Fixedwing aircraft must therefore achieve a minimum ground velocity ontakeoff before the lift on the wing is sufficient to overcome the weightof the plane. Fixed wing aircraft therefore generally require a longrunway along which to accelerate to achieve this minimum velocity andtakeoff.

In contrast, rotating wing aircraft can take off and land vertically oralong short runways inasmuch as powered rotation of the rotating wingprovides the needed lift. This makes rotating wing aircraft particularlyuseful for landing in urban locations or undeveloped areas without aproper runway.

The most common rotorcraft in use today are helicopters. A helicoptertypically includes an airframe or fuselage, housing an engine andpassenger compartment, and a rotor, driven by the engine, to providelift. Forced rotation of the rotor causes a reactive torque on thefuselage. Accordingly, conventional helicopters require either twocounter rotating rotors or a tail rotor in order to counteract thisreactive torque.

Another type of rotorcraft is the autogyro. An autogyro aircraft deriveslift from an unpowered, freely rotating rotor or plurality of rotaryblades. The energy to rotate the rotor results from a windmill-likeeffect of air passing through the underside of the rotor. The forwardmovement of the aircraft comes in response to a thrusting engine such asa motor driven propeller mounted fore or aft.

During the developing years of aviation aircraft, autogyro aircraft wereproposed to avoid the problem of aircraft stalling in flight and toreduce the need for runways. The relative airspeed of the rotating wingis independent of the forward airspeed of the autogyro, allowing slowground speed for takeoff and landing, and safety in slow-speed flight.Engines may be tractor-mounted on the front of an autogyro orpusher-mounted on the rear of the autogyro.

Airflow passing the rotary wing, alternately called rotor blades, whichare tilted upward toward the front of the autogyro, act somewhat like awindmill to provide the driving force to rotate the wing, i.e.autorotation of the rotor. The Bernoulli effect of the airflow movingover the rotor surface creates lift.

Various autogyro devices in the past have provided some means to beginrotation of the rotor prior to takeoff, thus further minimizing thetakeoff distance down a runway. One type of autogyro is the “gyrodyne,”which includes a gyrodyne built by Fairey aviation in 1962 and the XV-1convertiplane first flight tested in 1954. The gyrodyne includes athrust source providing thrust in a flight direction and a large rotorfor providing autorotating lift at cruising speeds. To provide initialrotation of the rotor, jet engines were secured to the tip of each bladeof the rotor and powered during takeoff, landing, and hovering.

Although rotorcraft provide the significant advantage of verticaltakeoff and landing (VTOL), they are much more limited in their maximumflight speed than are fixed wing aircraft. The primary reason that priorrotorcraft are unable to achieve high flight speed is a phenomenon knownas “retreating blade stall.” As the fuselage of the rotorcraft moves ina flight direction, rotation of the rotor causes each blade thereof tobe either “advancing” or “retreating.”

That is, in a fixed-wing aircraft, all wings move forward in fixedrelation, with the fuselage. In a rotary-wing aircraft, the fuselagemoves forward with respect to the air. However, rotor blades on bothsides move with respect to the fuselage. Thus, the velocity of any pointon any blade is the velocity of that point, with respect to thefuselage, plus the velocity of the fuselage. A blade is advancing if itis moving in the same direction as the flight direction. A blade isretreating if it is moving opposite the flight direction.

The rotor blades are airfoils that provide lift that depends on thespeed of air flow thereover. The advancing blade therefore experiencesmuch greater lift than the retreating blade. One technical solution tothis problem is that the blades of the rotors are allowed to “flap.”That is, the advancing blade is allowed to fly or flap upward inresponse to the increased air speed thereover such that its blade angleof attack is reduced. This reduces the lift exerted on the blade. Theretreating blade experiences less air speed and tends to fly or flapdownward such that its blade angle of attack is increased, whichincreases the lift exerted on the blade.

Flap enables rotating wing aircraft to travel in a directionperpendicular to the axis of rotation of the rotor. However, liftequalization due to flapping is limited by a phenomenon known as“retreating blade stall.” As noted above, flapping of the rotor bladesincreases the angle of attack of the retreating blade. However, atcertain higher speeds, the increase in the blade angle of attackrequired to equalize lift on the advancing and retreating blades resultsin loss of lift (stalling) of the retreating blade.

A second limit on the speed of rotorcraft is the drag at the tips of therotor. The tip of the advancing blade is moving at a speed equal to thespeed of the aircraft and relative to the air, plus the speed of the tipof the blade with respect to the aircraft. That is equal to the sum ofthe flight speed of the rotorcraft plus the product of the length of theblade and the angular velocity of the rotor. In helicopters, the rotoris forced to rotate in order to provide both upward lift and thrust inthe direction of flight. Increasing the speed of a helicopter thereforeincreases the air speed at the rotor or blade tip, both because of theincreased flight speed and the increased angular velocity of the rotorsrequired to provide supporting thrust.

The air speed over the tip of the advancing blade can therefore exceedthe speed of sound even though the flight speed is actually much less.As the air speed over the tip approaches the speed of sound, the drag onthe blade becomes greater than the engine can overcome. In autogyroaircraft, the tips of the advancing blades are also subject to thisincreased drag, even for flight speeds much lower than the speed ofsound. The tip speed for an autogyro is typically smaller than that of ahelicopter, for a given airspeed, since the rotor is not driven.Nevertheless, the same drag increase occurs eventually.

A third limit on the speed of the rotorcraft is reverse air flow overthe retreating blade. As noted above, the retreating blade is travelingopposite the flight direction with respect to the fuselage. At certainhigh speeds, portions of the retreating blade are moving rearward, withrespect to the fuselage, slower than the flight speed of the fuselage.Accordingly, the direction of air flow over these portions of theretreating blade is reversed from that typically designed to generatepositive lift. Air flow may instead generate a negative lift, ordownward force, on the retreating blade. For example, if the blade angleof attack is upward with respect to wind velocity, but wind is movingover the wing in a reverse direction, the blade may experience negativelift.

The ratio of the maximum air speed of a rotorcraft to the maximum airspeed of the tips of the rotor blades is known as the “advance ratio.The maximum advance ratio of rotorcraft available today is less than0.5, which generally limits the top flight speed of rotorcraft to lessthan 200 miles per hour (mph). For most helicopters, that maximumachievable advance ratio is between about 0.3 and 0.4.

In view of the foregoing, it would be an advancement in the art toprovide a rotating wing aircraft capable of vertical takeoff and landingand flight speeds in excess of 200 mph. It would also be an advance toprovide controls for the rotary wing that are comparatively stiffer,more precise, and containing less slack and backlash than prior artrotorcraft for timely responsiveness at such high speeds.

BRIEF SUMMARY OF THE INVENTION

The invention has been developed in response to the present state of theart and, in particular, in response to the problems and needs in the artthat have not yet been fully solved by currently available apparatus andmethods. The features and advantages of the invention will become morefully apparent from the following description and appended claims, ormay be learned by practice of the invention as set forth hereinafter.

A rotor system is disclosed including a mast having proximal and distalends. A rotor hub is rotatably mounted to the mast proximate the distalend. A plurality of rotor blades, extend from the hub, each having aproximal end rotatably mounted to the hub and a pitch control armsecured nearby. A swashplate encircles the mast and has a rotating plateand a non-rotating plate. A plurality of pitch control rods couple thepitch control arms to the non-rotating plate.

A plurality of swashplate actuators is rigidly mounted to the mast andcoupled to the non-rotating plate. The swashplate actuators areselectively activated to change an orientation and/or position of thenon-rotating plate. In some embodiments, a mast flange rigidly securedto, or monolithically formed with, the mast proximate the distal (e.g.,upper) end thereof and the swashplate actuators are rigidly mounted tothe mast flange.

The mast may be mounted to a pivot and a mast tilt actuator. The masttilt actuator and a mast pivot may be secured to the mast flange. Themast tilt actuator and pivot may be mounted to at least one vibrationsuppression component. In some embodiments, a shroud surrounds at leasta portion of the mast and defines a mast fluid path between the shroudand mast.

A lower edge of the shroud interfaces with the mast flange to form aseal operable to hinder airflow therethrough. An air plenum supplyingcompressed air during at least vertical operation, such as, takeoff,landing, and hover, is in fluid communication with the mast fluid path.The plenum may have an orientation with respect to the shroud effectiveto cause rotational flow of air forced into the mast fluid path to flowaround the mast from the plenum.

An upper edge of the shroud typically does at least one of eitherencircling a portion of the rotor hub and being encircled by the rotorhub. One or more rotary seals secured to at least one of the hub andshroud will operate to hinder airflow between the rotor hub and shroud.In some embodiments, the rotary seal includes a flexible skirt securedto an inner surface of one of the rotor hub and shroud.

The plurality of blades may each include a blade duct extending along atleast a portion of the length thereof. In such embodiments, the rotorhub defines a rotor cavity in fluid communication with the blade ductsof the plurality of blades and the mast fluid channel. The plurality ofblades each comprises a blade spar. The rotor hub comprises a pluralityof blade spar apertures, each aperture having a blade spar extendingthereinto. In such embodiments, the pitch control arms may be positionedwithin the rotor cavity.

The blade ducts may include hollow portions of the blade spars in fluidcommunication with the rotor cavity. Blade duct fairings may bepositioned within the rotor cavity around a distal edge of one of theblade ducts and have a contour effective to reduce pressure losses ofair flowing from the rotor cavity into the blade duct.

In some embodiments, a mast fairing is secured to the rotor hub andencircles the mast along a portion thereof between the proximal anddistal ends thereof. The mast fairing may have a contour selected toreduce pressure losses of air flowing from the mast fluid path into therotor cavity.

BRIEF DESCRIPTION OF THE DRAWINGS

The foregoing features of the present invention will become more fullyapparent from the following description and appended claims, taken inconjunction with the accompanying drawings. Understanding that thesedrawings depict only typical embodiments of the invention and are,therefore, not to be considered limiting of its scope, the inventionwill be described with additional specificity and detail through use ofthe accompanying drawings in which:

FIG. 1 is an isometric view of an aircraft in accordance with anembodiment of the present invention;

FIG. 2 is a front elevation view of a compressed or otherwisepressurized air supply for a tip jet in accordance with an embodiment ofthe present invention;

FIG. 3A is a front elevation view of a rotorcraft illustratingoperational parameters describing a rotor configuration suitable for usein embodiments of an apparatus and method in accordance with the presentinvention, and the system of FIGS. 1 and 2 in particular;

FIG. 3B is a right side elevation view of the rotorcraft of FIG. 3A;

FIG. 3C is a partial cut of a right side elevation view of the rotor ofFIG. 3A;

FIG. 4A is a side elevation view of an embodiment of a rotor inaccordance with the present invention;

FIG. 4B is a side elevation cross-sectional view of an embodiment of arotor in accordance with the invention;

FIG. 4C is a bottom plan view of an embodiment of a rotor in accordancewith the present invention;

FIG. 5 is a partial isometric view of an embodiment of a rotor inaccordance with the present invention;

FIG. 6 is a side elevation cross-sectional view of an embodiment of arotor and mounting system in accordance with the present invention;

FIG. 7 is a top cross-sectional view of an embodiment of a rotor havingan angled air inlet in accordance with the present invention;

FIG. 8 is a side elevation cross-sectional view of an embodiment of ahub-shroud seal for a rotor in accordance with the present invention;

FIG. 9 is a side elevation cross-sectional view of an embodiment of ashroud-mast flange seal for a rotor in accordance with the presentinvention;

FIG. 10 is a side elevation cross-sectional view of an embodiment of aswashplate for a rotor in accordance with the present invention;

FIG. 11 is an isometric view of a rotating ring of an embodiment of aswashplate for a rotor in accordance with the present invention;

FIG. 12 is an isometric view of a non-rotating ring of an embodiment ofa swashplate for a rotor in accordance with the present invention;

FIG. 13A is a bottom plan view of an embodiment of a non-rotating ringhaving oil ports for a swashplate for use in a rotor in accordance withthe present invention;

FIG. 13B is a side elevation cross-sectional view of an embodiment of anoil port of the non-rotating ring of a swashplate for use in a rotor inaccordance with the present invention;

FIG. 14A is an isometric view of an upper seal seat of an embodiment ofa swashplate for a rotor in accordance with the present invention;

FIG. 14B is an isometric view of a lower seat of an embodiment of aswashplate for a rotor in accordance with the present invention;

FIG. 15 is a partial side elevation cross-sectional view of anembodiment of a swashplate for a rotor in accordance with the presentinvention;

FIG. 16 is a side elevation cross-sectional view of an embodiment of amast bearing for a rotor in accordance with the present invention;

FIG. 17 is a side elevation cross-sectional view of an embodiment of arotor hub for a rotor in accordance with the present invention;

FIG. 18 is another side elevation cross-sectional view of an embodimentof a mast and hub assembly for a rotor in accordance with the presentinvention;

FIG. 19 is a side elevation cross-sectional view of an alternativeembodiment of a mast and hub assembly for a rotor in accordance with anembodiment of the present invention;

FIG. 20 is a partial isometric cross-sectional view of an embodiment ofa seal mounting ring for a rotor in accordance with the presentinvention;

FIG. 21 is a side elevation cross-sectional view of an embodiment of aspindle bearing for a rotor in accordance with the present invention;

FIG. 22 is a schematic diagram of an embodiment of an oil distributionsystem for a rotor in accordance with the present invention;

FIG. 23A is a schematic diagram of an embodiment of a passive air flowheating system for a rotor in accordance with the present invention;

FIG. 23B is a schematic diagram of an embodiment of a bleed air heatingsystem for a rotor in accordance with the present invention;

FIG. 24 is a block diagram of an embodiment of a flight control systemfor an aircraft incorporating a rotor in accordance with the presentinvention;

FIG. 25 is a process flow diagram of an embodiment of a method forthermal management of a rotor in accordance with the present invention;

FIG. 26 is a process flow diagram of an embodiment of another method forextracting heat from a rotor in accordance with the present invention.

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS

It will be readily understood that the components of the presentinvention, as generally described and illustrated in the drawingsherein, could be arranged and designed in a wide variety of differentconfigurations. Thus, the following more detailed description of theembodiments of the system and method of the present invention, asrepresented in the drawings, is not intended to limit the scope of theinvention, as claimed, but is merely representative of variousembodiments of the invention. The illustrated embodiments of theinvention will be best understood by reference to the drawings, whereinlike parts are designated by like numerals throughout.

Referring to FIG. 1, an aircraft 10 includes a fuselage 12 or airframe12 defining a cabin for carrying an operator, passengers, cargo, or thelike. The airframe 12 may include one or more fixed wings 14 shaped asairfoils for providing lift to the aircraft. The wings 14 may beconfigured such that they provide sufficient lift to overcome the weightof the aircraft 10 only at comparatively high speeds inasmuch as theaircraft 10 is capable of vertical takeoff and landing (VTOL) and doesnot need lift from the fixed wings 14 at low speeds, e.g. below 50 mphor even 100 mph upon taking off.

In this manner, the wings 14 may be made smaller than those of fixedwing aircraft requiring a high velocity takeoff, which results in lowerdrag at higher velocities. In some embodiments the wings 14 providesufficient lift to support at least 50 percent, preferably 90 percent,of the weight of the aircraft 10 at air speeds above 200 mph.

Control surfaces 16 may secure to one or both of the airframe 12 andwings 14. For example a tail structure 18 may include one or morevertical stabilizers 20 and one or more rudders 22. The rudders 22 maybe adjustable as known in the art to control the yaw 24 of the aircraft10 during flight. As known in the art, yaw 24 is defined as rotationabout a vertical axis 26 of the aircraft 10. In the illustratedembodiment, the rudders 22 may comprise hinged portions of the verticalstabilizers 20.

The tail structure 18 may further include a horizontal stabilizer 28 andan elevator 30. The elevator 30 may be adjustable as known in the art toalter the pitch 32 of the aircraft 10. As known in the art, pitch 32 isdefined as rotation in a plane containing the vertical axis 26 and alongitudinal axis 34 of the airframe of an aircraft 10. In theillustrated embodiment, the elevator 30 is a hinged portion of thehorizontal stabilizer 28. In some embodiments, twin rudders 22 may bepositioned at an angle relative to the vertical axis 26 and serve bothto adjust the yaw 24 and pitch 32 of the aircraft 10.

The control surfaces 16 may also include ailerons 36 on the wings 14. Asknown in the art, ailerons 36 are used to control roll 38 of theairplane. As known in the art, roll 38 is defined as rotation about thelongitudinal axis 34 of the aircraft 10.

Lift during vertical takeoff and landing and for augmenting lift of thewings 14 during flight is provided by a rotor 40 comprising a number ofindividual blades 42. The blades are mounted to a rotor hub 44. The hub44 is coupled to a mast 46 which couples the rotor hub 44 to theairframe 12. The rotor 40 may be selectively powered by one or moreengines 48 housed in the airframe 12, or adjacent nacelles, and coupledto the rotor 40. In some embodiments, jets 50 located at or near thetips of the blades 42 power the rotor 40 during takeoff, landing,hovering, or when the flight speed of the aircraft is insufficient toprovide sufficient autorotation to develop needed lift.

Referring to FIG. 2, while still referring to FIG. 1, in the illustratedembodiment, the engines 48 may be embodied as jet engines 48 thatprovide thrust during flight of the aircraft. The jet engines 48 mayadditionally supply compressed air to the jets 46 by driving a bypassturbine 62 or auxiliary compressor. Air compressed by the bypass turbine62 may be transmitted through ducts 54 to a plenum 56 in fluidcommunication with the ducts 54.

The plenum 56 is in fluid communication with the mast 46 that is hollowor has another passage to provide for air conduction. A mast shroud 58positioned around the mast 46 may provide one or both of an air channeland a low drag profile for the mast 46. The mast 46 or mast shroud 58 isin fluid communication with the rotor hub 44. The rotor hub 44 is influid communication with blade ducts 60 extending longitudinally throughthe blades 42 to feed the tip jets 50.

Referring to FIGS. 3A-3C, rotation of the rotor 40 about its axis ofrotation 72 occurs in a rotor disc 70 that is generally planar but maybe contoured due to flexing of the blades 42 during flight. In general,the rotor disc 70 may be defined as a plane in which the tips of theblades 42 travel. Inasmuch as the blades 42 flap cyclically upward anddownward due to changes in lift while advancing and retreating, therotor disc 70 is angled with respect to the axis of rotation 72 whenviewed along the longitudinal axis 34, as shown in FIG. 3A.

Referring to FIG. 3B, the angle 74 of the rotor disc 70, relative to aflight direction 76 in the plane containing the longitudinal axis 34 andvertical axis 26, is defined as the angle of attack 74 or rotor diskangle of attack 74. For purposes of this application, flight direction76 and air speed refer to the direction and speed, respectively, of theairframe 12 of the aircraft 10 relative to surrounding air. In autogyrosystems, the angle of attack 74 of the rotor disc 70 is generallypositive in order to achieve autorotation of the rotor 40, which in turngenerates lift.

Referring to FIG. 3C, the surfaces of the blades 42, and particularlythe chord of each blade 42, define a pitch angle 78, or blade angle ofattack 78, relative to the direction of movement 80 of the blades 42. Ingeneral, a higher pitch angle 78 will result in more lift and higherdrag on the blade up to the point where stalling occurs, at which pointlift has declined below a value necessary to sustain flight. The pitchangle 78 of the blade 42 may be controlled by both cyclic and collectivepitch control as known in the art of rotary wing aircraft design.

Referring to FIGS. 4A through 4C, the mast shroud 58 encircles the mast46 such that the shroud and mast define an annular air channel 92. In analternative embodiment, the air channel 92 passes through the center ofthe mast 46. The hub 44 defines a cavity 94 in fluid communication withthe air channel 92. The blade ducts 60 are in fluid communication withthe cavity 94 enabling air flow from the air channel 92 through thecavity 94 and blade ducts 60 to the tip jets 50. In the illustratedembodiment, the blades 42 define a hollow blade spar 96 extendingthrough the hub 44 into the cavity 94 and the blade duct 60 is embodiedas a hollow channel extending longitudinally through the blade spar 96.

A mast flange 98 may be rigidly secured to, or formed monolithicallywith the mast 46. The shroud may create a substantially continuousbarrier to air flow between the hub 44 and the mast flange 46, but foran inlet 100 coupled to the plenum 56. By substantially continuousbarrier to air flow, what is meant is that the shroud ensures that atleast 90%, preferably at least 95%, of all air entering the air channel92 from the plenum 56 passes into the cavity 94 of the hub 44. Theshroud may additionally include one or more sealed hatches 102 that areselectively openable to service internal components of the rotor 40without requiring removal of the entire mast shroud 58.

Referring specifically to FIG. 4B, the one or more hatches 102 may belocated adjacent a swashplate 104 that encircles the mast 46 tofacilitate servicing of the swashplate 104. The swashplate 104 mayengage a spherical bearing 106 or gimbal slidingly mounted to the mast46. The swashplate 104 includes a rotating ring 108 and a non-rotatingring 110 rotatably secured to one another. The swashplate 104 mayencircle the mast 46 between the hub 44 and the mast flange 98. Thenon-rotating ring 110 is coupled to swashplate actuators 112, such as bymeans of actuator rods 114. The rotating ring 108 is coupled to theblade spars 96, such as by means of pitch horns 118 coupled nearproximal ends 120 of the blade spars 96 and pitch control rods 116connecting the pitch horns 118 to the rotating ring 108. The swashplateactuators 112 are selectively actuated to raise and lower the swashplate104 and spherical bearing 106 in order to change the pitch angles 78 ofall of the blades 42 by a uniform amount, i.e., the collective pitch ofthe blades 42. The swashplate actuators 112 are also selectivelyactuated to change the angle of the swashplate 104 in order to changethe amplitude and phase of cyclic variation of the pitch angles 78,i.e., cyclic pitch, of the blades 42 as they rotate around the mast 46.The swashplate actuators 112 may be selectively activated to change thecollective and cyclic pitch simultaneously.

The swashplate actuators 112 are rigidly mounted to the mast 46, such asby rigidly mounting the swashplate actuators 112 to the mast flange 98.In the illustrated embodiment, the swashplate actuators 112 are securedon an opposite side of the flange 98 as the swashplate 104 and the mastflange 98 defines apertures 122 permitting the actuator rods 114 to passfrom the swashplate actuator 112 to the swashplate 104. The swashplateactuators 112 may be embodied as hydraulic pistons 124 and cylinders126.

Referring specifically to FIG. 4C, in the illustrated embodiment, threeswashplate actuators 112 secure to the flange 98 inasmuch as threeindependent points are sufficient to define any plane, i.e., anyorientation of the swashplate 104. The swashplates 112 may bedistributed at equal or unequal angular intervals around the axis ofrotation 72 of the hub 44 and may be located at equal or unequaldistances from the axis of rotation of the hub 44.

Referring again to FIG. 4B, a central channel 128 may extend throughboth the mast 46 and the mast flange 98. A plurality of electrical lines130 and fluid lines 132, such as oil and fuel lines, may pass throughthe central channel 128. The lines 130, 132 may couple to the hub 44 orstructures fixedly mounted to the hub 44. Accordingly, the lines 130,132 may rotate with the hub 44. The electrical lines 130 may couple to aslip ring assembly 134 coupling signals from stationary lines 136 to therotating electrical lines 130. In a like manner, a hydraulic rotaryunion 138 may couple stationary fluid ports 140 to the rotating fluidlines 132.

The air channel 92 enables the flow of air from the inlet 100 to thecavity 94 and into the blade ducts 60. In some embodiments, structuresof the rotor 40 include fairings to reduce drag on moving components andpressure losses incurred on air moving from the inlet 100 to the bladeducts. For example, a mast fairing 142 may secure to the hub 44 andencircle that mast 46. The mast fairing 142 extends along the mast 46and has a contour effective to reduce pressure losses in air flowingfrom along the mast 46 to along the hub 44. For example, the mastfairing may have an outer surface 144 that decreases smoothly indiameter with distance from the hub 44 along the axis of rotation 72. Bysmoothly, what is meant is that the slope of the change in outerdiameter with distance from the hub 44 along the axis of rotation 72does not exceed 1.0, except for possibly discontinuities at the upperand lower edges of the mast fairing 142. In some embodiments, the mastfairing 142 is fixed to the mast and the hub 44 is free to rotaterelative to the mast fairing 142.

Pressure losses in airflow from the cavity 94 to the blade ducts 60 mayalso be reduced by means of a blade duct fairing 146 covering theproximal end 120 of the blade spar 96. The blade duct fairing 146defines a “bell mouth” as known in the art of aerodynamics. The bladeduct fairing 146 may define an aperture 148 through which the pitchhorns 118 protrude.

FIG. 5 illustrates an isometric view of the swashplate 104. In additionto the pitch control rods 116 and actuator rods 114 coupled to therotating and non-rotating rings 108, 110, respectively, rotatingstabilizer linkages 150 and non-rotating stabilizer linkages 152 maysecure to the rotating and non-rotating rings 108, 110, respectively.The linkage 150 hinders rotation of the rotating ring 108 relative tothe hub 44 while still permitting vertical movement of the rotating ring108 relative to the hub 44. Likewise, the linkage 152 hinders rotationof the non-rotating ring 110 relative to the swashplate actuators 112.The linkages 150, 152 may each comprise a leg 154 a coupled to therotating ring 108 and non-rotating ring 110, respectively, and a leg 154b coupled to the hub 44 and mast 46, respectively. In the illustratedembodiment, the leg 154 b of the linkage 150 is mounted to the mastfairing 142 secured to the hub 44 and the leg 154 b of the linkage 152is mounted to the mast flange 98.

The legs 154 a, 154 b are coupled to one another by a hinge 156 defininga pivoting axis 158 that is perpendicular to the axis of rotation 72 ofthe hub 44 and tangent to a circle centered on the axis of rotation 72of the hub 44. The linkages 150, 152 reduce torque on the pitch controlrods 116 and actuator rods 114 which are oriented vertically and are notwell suited to bear such loads without hindering their ability to movefreely in response to actuator inputs.

The rotating ring 108 may have one or more linkage mounts 160 securedthereto, or formed monolithically therewith. The legs 154 a of one ormore linkages 150 may secure to the linkage mounts 160, such as by meansof a spherical joint. The non-rotating ring may likewise have one ormore linkage mounts 162 secured thereto, or formed monolithicallytherewith. The legs 154 a of one or more linkages 152 may secure to thelinkage mounts 162, such as by means of a spherical joint.

Referring to FIG. 6, the mast flange 98 may include a pivot 170pivotally secured to a mount 172 enabling changing of the angle ofattach 74 of the rotor disc 70. An actuator 174 may secure to a mount172 and flange 98 for changing the pitch of the hub 44 and 46. Theactuator 174 may be embodied as a hydraulic cylinder 176 and piston 178having one of the cylinder 176 and piston 178 coupled to the flange 98and the other of the hydraulic cylinder 176 and piston 178 coupled tothe mount 172. The actuator 174 and pivot 170 may secure to separatemounts 172 or may be attached to the same mount 172. The one or moremounts 172 are secured to a structural member 180 of the fuselage 12 bymeans of one or more vibration suppression devices 182 or dampers 182.

The above described arrangement of the pivot 170, actuator 174, mount172, and vibration suppression 182 in conjunction with the mounting ofthe swashplate actuators 112 to the mast 46, such as the mast flange 98,provide a rotor 40 that is exceptionally rigid with very little play orslop between the blades 42 and the mast 46 and between the pitch horns118 and the swashplate actuators 112. Due to the rigidity of the rotor40, the frequency response of the rotor 40 in the same range offrequencies as the cyclic loads on the rotor 40 induced by cyclicvariation in lift on the blades 42 may be damped, rather than resonant.For high speed flight, the frequency of rotation of the blades 42 isgenerally reduced to a minimum frequency of rotation in order to avoidthe problems mentioned hereinabove that occur at high advance ratiossuch as retreating blade stall and high tip speed mach numbers.Reduction of rotation frequency may also facilitate equalization of rollmoments exerted by the blades 42 as described in U.S. Prov. Pat. App.Ser. No. 61/403,136, filed Sep. 9, 2010, and entitled ROLL MOMENTEQUALIZATION AT HIGH ADVANCE RATIOS. Accordingly, low frequency cyclicloading may occur with large amplitudes during high speed flight. Therigidity of the rotor system described hereinabove, increases theharmonic frequencies of the rotor in order to reduce the risk ofdestructive resonance for low rotational frequencies during high speedflight. The rigidity of the coupling between the swashplate actuators112 and the pitch horns 118 likewise enables very precise control ofcollective and cyclic pitch and raises harmonic frequencies of thelinkage between the swashplate actuators 112 and the pitch horns 118 inorder to avoid destructive resonances during high speed flight.

Referring to FIG. 7, in some embodiments, the inlet 100 defines a centeraxis 190 having a non-zero angle 192 with respect to the vertical plane196 containing the vertical axis 26 and longitudinal axis 34 of theaircraft 10. In some embodiments the angle is between 5 and 45 degrees,preferably between 10 and 30 degrees. In some embodiments, the extensionof the center axis 190 of the inlet across the air channel 92 is offseta distance 198 from the axis of rotation 72 of the hub 44 at its pointof greatest proximity to the axis of rotation 72. The distance 190 maybe between may be between 5 and 40 percent, preferably between 15 and 35percent of the diameter of the shroud 58. In some embodiments, one orboth of the angle 192 and offset distance 190 may be chosen effective toreduce the pressure losses of air flowing through the air channel 92 andcavity 94 by between five percent and fifty percent, preferably betweentwenty and fifty percent.

As a result of the angle 192 and/or offset distance 198, air flow 196within the inlet 100 will be forced to rotate within the annular airchannel 92 defined between the mast 46 and the mast shroud 58. Duringoperation, the rotor 40 includes a number of rotating componentsincluding the pitch control rods 116, rotating ring 108, stabilizerlinkages hub 44, and pitch horns 118, that are rotating at high speeds.Due to the rotational velocity of the air within the inlet, the relativeair speed between the air flow 196 and the rotating components of therotor 40 is reduced, provided the rotational velocity of the air flow196 is in the same direction as the tangential velocity of the rotatingcomponents of the rotor 40. Accordingly, drag on the rotating componentsand pressure losses of the air flow 196 over the rotating components,which are proportional to the relative velocity squared, will bereduced. Pressure losses of the air flow 196 will also be reducedinasmuch as the air entering the air channel 92 from the inlet 100 isnot required to make a 90 degree turn, which is very aerodynamicallyinefficient. In some embodiments, one or both of the angle 192 andoffset distance 190 may be chosen effective to reduce pressure losses ofair flowing through the air channel 92 and cavity 94 by between fivepercent and twenty percent, preferably between ten percent and thirtypercent.

In some embodiments, the inlet 100 may have curved walls such that acenter axis is not readily identified. In such embodiments, and inembodiments having a generally straight inlet 100, the contour andorientation of the inlet 100 may be such that air from the inlet 100flowing through the air channel 92 while the tip jets 50 are ignited hasan average angular velocity that is in the same direction as the angularvelocity of the hub and has a magnitude greater than 50 percent,preferably greater than 80 percent, of the magnitude of the angularvelocity of the hub 44.

Referring to FIG. 8, in order to reduce pressure losses due to airleakage, a seal 200 may be secured to one of the hub 44 and shroud 58.The seal 202 may include a ring 202 of flexible material secured to oneof the hub 44 and mast shroud 58 and overlapping a portion of the otherof the hub 44 and the mast shroud 58. In the illustrated embodiment, thering 202 secures to an outer surface 204 of the mast shroud 58 andoverlaps an inner surface 206 of the hub 44. When pressurized air iswithin the air channel 92, the ring 202 is urged against the innersurface 206 of the hub 44 and hinders air leakage.

In some embodiments, the ring 202 is an inner ring 202 and the seal 200includes an outer ring 208. The inner and outer rings 202, 208 mayinclude cuts 210 enabling the rings 202, 208 to flair outwardly inresponse to air pressure within the shroud 58. The cuts 210 of the innerring 202 may be offset from the cuts 210 of the outer ring 208 such thatair flow through aligned cuts 210 is prevented. The rings 202, 208 maybe made of a flexible polymer with high wear resistance or coated with awear resistant material. For example, the rings 202, 208 may be made ofwear resistant polymer.

Referring to FIG. 9, air leakage between the mast shroud 58 and the mast46 may be hindered by a flange 220 secured to, or formed monolithicallywith, the mast shroud 58 and encircling the mast shroud 58. The flange220 is secured to a seat 222 formed on the mast flange 98. A seal orsealant material may be interposed between the flange 220 and seat 222.Alternatively, mating surfaces of the flange 220 and seat 222 may bepolished sufficiently to create an adequate seal. Fasteners 224, such asscrews or bolts, distributed circumferentially around the flange 200 mayengage the flange 220 and seat 222 to secure the flange 220 to the seat222 and promote creation of a seal therebetween. In an alternativeembodiment, a flange secured to, or formed monolithically with, the mastflange and projecting upwardly from the mast flange 98 may secure to theshroud 58 and the seat 22 and flange 220 may be omitted.

FIG. 10 is a side elevation cross-sectional view of a swashplate 104. Asnoted above, the swashplate 104 includes a rotating ring 108 and anon-rotating ring 110. The non-rotating ring 110 is mounted to aspherical bearing 106 that is slidably mounted to the mast 46. In theillustrated embodiment, the rotating ring 108 includes pitch control rodmounts 230 secured to or monolithically formed therewith and pivotallysecured to the pitch control rods 116. Likewise, the non-rotating ring110 includes actuator rod mounts 232 secured there to or monolithicallyformed therewith. The actuator rods 114 pivotally secure to the actuatorrod mounts 232.

An upper seal 234 and a lower seal 236 are interposed between therotating ring 108 and the non-rotating ring 110. One or two bearings238, 240 may likewise be interposed between the rotating ring 108 andthe non-rotating ring 110. The seals 234, 236 may be positioned withinseal seats 242, 244, respectively. The seats 242, 244 may be embodied asseparate members secured to the non-rotating ring 110 and rotating ring108, respectively. In the illustrated embodiments, the seat 242 alsocapture a portion of the bearings 238, 240 between itself and thenon-rotating ring 110. In a like manner, the seat 244 captures a portionof the bearings 238, 240 between itself and the rotating ring 108.

An upper clamping ring 246 captures the seal 234 between itself and theseat 242. Likewise, a lower clamping ring 248 captures the seal 236between itself and the seat 244. The upper clamping ring 246 may besecured to the seat 242 by means of fasteners 250, such as screws,bolts, or the like. The fasteners 250 may extend through both the upperclamping ring 246 and the seat 242 and fasten to the non-rotating ring110 thereby securing both the clamping ring 246 and the seat 242 to thenon-rotating ring 110. The upper clamping ring 246 may additionallycapture the spherical bearing 106 between itself and the non-rotatingring 110. The upper clamping ring 246 may define a spherical bearingseat 252 having a spherical contour for engaging the spherical bearing106. In a like manner, fasteners 254 may secure the lower clamping ring248 to the seat 244 and may pass through both the lower clamping ring248 and the seat 244 and secure to the rotating ring 108, therebysecuring both the lower clamping ring 248 and the seat 244 to therotating ring 108. A sealing material, such as a polymer gasket, may bepositioned between the upper clamping ring 246 and the seat 242 andbetween the lower clamping ring 248 and the seat 244 to create a sealtherebetween hindering the leakage of oil.

Referring to FIG. 11, the rotating ring 108 defines a bearing seat 260sized to receive a portion of the bearings 238, 240. The bearing seat260 may be embodied as an inward facing cylindrical wall 262 and aradial wall 264 extending radially inward from one edge of thecylindrical wall 262. The bearings 238, 240 may be captured within thebearing seat 260 by means of the seal seat 244 secured to the rotatingring 108. The rotating ring 108 may additionally include a sealingsurface 266 positioned to engage the upper seal 234. In the illustratedembodiment, the sealing surface 266 is a cylindrical surface centered onthe central axis 268 of the rotating ring 108. The cylindrical wall 262may likewise be centered on the central axis 268.

The bearing seat 260 may have a plurality of grooves formed therein tofacilitate the flow of oil around and through the bearings 238, 240. Thegrooves may include circumferential grooves 270 formed in one or both ofthe cylindrical wall 262 and the radial wall 264. The grooves may alsoinclude grooves 272 extending vertically along the cylindrical wall 262and radially along the radial wall 264. The rotating ring 108 mayadditionally define apertures 274, which may be threaded, for receivingthe fasteners 254.

Referring to FIG. 12, the non-rotating ring 110 defines a bearing seat280 sized to receive a portion of the bearings 238, 240. The bearingseat 280 may be embodied as an outward facing cylindrical wall 282 and aradial wall 284 extending radially outward from one edge of thecylindrical wall 282. The bearings 238, 240 may be captured within thebearing seat 280 by means of the seal seat 242 secured to thenon-rotating ring 110. The non-rotating ring 110 may additionallyinclude a sealing surface 286 positioned to engage the lower seal 236.In the illustrated embodiment, the sealing surface 286 is a cylindricalsurface centered on the central axis 288 of the non-rotating ring 110.The cylindrical wall 282 may likewise be centered on the central axis288.

The bearing seat 280 may have a plurality of grooves formed therein tofacilitate the flow of oil around and through the bearings 238, 240. Thegrooves may include circumferential grooves 290 formed in one or both ofthe cylindrical wall 282 and the radial wall 284. The grooves may alsoinclude grooves 292 extending vertically along the cylindrical wall 282and radially along the radial wall 284.

The non-rotating ring 110 may additionally define apertures 294, whichmay be threaded, for receiving the fasteners 250. The non-rotating ring110 may also define a spherical bearing surface 296 engaging thespherical bearing 106 in the assembled swashplate 104. The sphericalbearing surface 296 may have a spherical contour sized to mate with thespherical bearing 106.

Referring to FIGS. 13A and 13B, the non-rotating ring 110 may include afeed port 300 a and a return port 300 b mounted thereto. Each of thefeed port 300 a and return port 300 b is in fluid communication with achannel 302 in fluid communication with one or both of a vertical groove292 and a circumferential groove 290. Oil for lubricating the bearings238, 240 is pumped into the feed port 300 a and is forced out of thereturn port 300 b. In addition to lubrication, the oil may be used tocool or heat the bearings 238, 240 in order to maintain the bearings238, 240 in a preloaded condition notwithstanding temperature variationsof air flow over the swashplate 104 and heat buildup due to friction. Inan alternative embodiment, one or both of the feed port 300 a and thereturn port 300 b are secured to the rotating plate 108. In suchembodiments, the feed port 300 a and/or return port 300 b coupled to therotating plate 108 may be connected to fluid lines 132 emanating fromthe hydraulic rotary union 138.

Referring to FIG. 14A, the seal seat 242 may be embodied as a ringincluding radial grooves 310 formed on a lower surface thereof andextending partially radially inwardly from an outer circumference of thelower surface. The grooves 310 may be positioned such that the grooves310 are aligned with the vertical grooves 292 of the non-rotating ring110 when in place as shown in FIG. 10. The seal seat 242 may likewiseinclude apertures 312 for receiving the fasteners 250.

The seal seat 242 may include an outwardly facing cylindrical wall 314centered on the central axis 316 of the seal seat 242 and a radial wall318 extending radially outward from the central axis 316 from an edge ofthe cylindrical wall 314. The seal 234 may abut the cylindrical wall 314and the radial wall 318 in the assembled swashplate 104.

Referring to FIG. 14B, in a like manner, the seal seat 244 may beembodied as a ring including radial grooves 320 formed on an uppersurface thereof and extending partially radially outwardly from an innercircumference of the upper surface. The grooves 320 may be positionedsuch that the grooves 320 are aligned with the vertical grooves 272 ofthe rotating ring 108 when in place as shown in FIG. 10. The seal seat244 may likewise include apertures 322 for receiving the fasteners 254.

The seal seat 244 may include an inwardly facing cylindrical wall 324centered on the central axis 326 of the seal seat 244 and a radial wall328 extending radially inward from the central axis 326 from an edge ofthe cylindrical wall 324. The seal 236 may abut the cylindrical wall 324and the radial wall 328 in the assembled swashplate 104.

Referring to FIG. 15, the radial grooves 310 of the seal seat 242, thegrooves 272 of the rotating plate 108, the grooves 292 of thenon-rotating plate 110, and the radial grooves 320 of the seal seat 244form a fluid path extending around the bearings 238, 240. The sealingsurface 266 of the rotating ring 108 engages the seal 234 and thesealing surface 286 of the non-rotating ring 110 engages the seal 236.The seals 234, 236 hinder the entry of contaminants into the bearings238, 240 and hinder the leakage of oil from between the rotating plate108 and the non-rotating plate 110.

The bearings 238, 240 may each include an outer race 330 engaging therotating ring 108 and moving synchronously therewith and an inner race332 engaging the non-rotating ring 110 and being fixed relative to thenon-rotating ring 110. A plurality of rolling elements 334, such as ballbearings, are captured between the outer race 330 and the inner race332. A cage 336 may also be positioned between the inner and outer races330, 332 to maintain the rolling elements separated from one another andevenly distributed around the races 330, 332.

The rolling elements 334 may be preloaded such that they are deformedfrom an undeformed shape even in the absence of any loads on therotating ring 108 or non-rotating ring 110. Preloading the rollingelements 334 may eliminate slop or play between the rotating ring 108and non-rotating ring 110 that would exist if gaps were present betweenthe rolling elements 334 and the inner and outer races 330, 332. Due tothermal contraction of the rolling elements 332, 334, rotating ring 108,and non-rotating ring 110, the preloaded condition of the rollingelements 334 may be reduced or disappear.

In embodiments of the present invention, air directed through the airchannel 92 to the tip jets 50 may be at an elevated temperature due tothe input of energy during compression of the air. In some embodiments,the temperature of air forced through the air channel 92 may be above300° F. when the tip jets 50 are ignited. As a result, during verticaltakeoff and landing or during hover, the preload of the bearing elements334 will be increased due to thermal expansion of the bearings elements334, races 330, 332, rotating ring 108, and non-rotating ring 110.However, during sustained longitudinal flight at high speeds andaltitudes, hot compressed is no longer needs to flow to the tip jets 50and the ambient air temperature can be very low. For example, above analtitude of 8000 ft, the air temperature is typically at or below 32° F.Accordingly, the preload of the rolling elements 334 may decrease tozero and gaps may occur between the rolling elements 334 and the races330, 332, resulting in increased slop or play between the rotating plate108 and the non-rotating plate 110. Increases in slop or play betweenthe rotating plate 108 and the non-rotating plate 110 may result indestructive harmonics at the frequency of cyclic loads on the blades 42during high speed flight.

Accordingly, in some embodiments, oil flowing through the bearings 238,240 may be selectively cooled to prevent over loading or heat relatedfailure of the rolling elements 334 due to hot air flow and heat buildupdue to friction. The oil flowing through the bearings 238, 240 may alsobe selectively heated to prevent cooling to the point that the rollingelements 334 are no longer preloaded or the preload of the rollingelements 334 is below a predetermined threshold.

Each of the seals 234, 236 may include an outer seal 340 and an innerseal 342. In the illustrated embodiment, the outer seal 340 and innerseal 342 are mirror images of one another. The outer seal 340 may beseparated from the inner seal 342 by a spacer 344. The seals 340, 342may include a sealing material 346 disposed in a ring and defining asealing surface 348 for engaging corresponding sealing surfaces 266, 286of the rotating ring 108 and non-rotating ring 110, respectively. Thesealing material 346 may define a groove 350 having a circumferentialspring 352 positioned therein and biased to urge the sealing surface 348against the sealing surface 266 or sealing surface 286. The groove 350of the upper seal 340 and the groove 350 of the inner seal 342 may faceaway from one another. The sealing material 346 may be mounted within aretainer 354 formed of a metal or other rigid material for maintainingthe shape of the sealing material 346 during use and installation of theupper and lower seals 340, 342.

The spacer 344 may include one or more grooves 356 extending radiallytherethrough. The grooves 356 may permit the passage of any oil leakingbetween the outer and inner seals 340, 342 to flow into a fluid path 358formed in the upper seal seat 242 non-rotating ring 110. The fluid path358 may be in fluid communication with the return port 300 b. The fluidpath 358 preferably connects to the fluid path between the feed port 300a and the return port 300 b at a point that is at a lower pressure thanoil flowing adjacent the inner seal 342 such that oil tends to flow onlyoutwardly from the space between the seals 340, 342 into the fluid path358.

Referring to FIG. 16, the hub 44 may mount to the mast 46 by means of anupper bearing 370 and a lower bearing 372 captured between the hub 44and the mast 46. The bearings 370, 372 may be embodied as tapered rollerbearings including tapered rolling elements 374. Tapered rollingelements 374 advantageously support loads perpendicular to the axis ofrotation 72 of the hub 44 and support lift loads parallel to the axis ofrotation 72. The axes 376 of the tapered rolling elements 376 form anangle 378 with the axis of rotation 72. The angle 378 is dictated by themagnitude of radial and longitudinal forces. For example, the angle 378may be between 3 and 20 degrees. The angle 378 of the rolling elements374 of the upper bearing 270 may be unequal that of the rolling elements374 of the lower bearing 372. As is apparent from FIG. 16, the rollingelements 374 of the upper bearing 370 are angled away from the axis ofrotation 72 with upward distance along the axis of rotation 72 whereasthe rolling elements 374 of the lower bearing 372 are angled away fromthe axis of rotation 72 with downward distance along the axis ofrotation 72, where the downward and upward directions are in the frameof reference of the page. The opposing orientation of the angles 378enables support of loads in both directions parallel to the axis ofrotation 72.

Each of the upper bearings 370, 372 includes a cup 380 and a cone 382 asknown in the art of tapered rolling design. The cup 380 extends aroundthe rolling elements 374 and includes a shallow channel 384, ordepression 384, for retaining the rolling elements 374. The cup 380 ofthe upper bearing 370 faces opposite the cup 380 of the lower bearing372. The cone 382 is located among the rolling elements 374 having therolling elements 374 captured between the cone 382 and the cup 384. Thecone 382 includes a channel 386, or depression 386, for retaining therolling elements 374. For each of the upper bearings 370, 372, therolling elements 374 are captured between the channel 384 of the cup 380and the channel 386 of the cone 382.

Referring to FIG. 17, the hub 44 includes an upper bearing seat 390 anda lower bearing seat 392. The cups 380 of the upper and lower bearings370, 372 are positioned within the bearings seats 390, 392,respectively. Each of the seats 390, 392 includes a vertical wall 394having a cylindrical shape parallel to the axis of rotation 72 of thehub 44. The seats 390, 392 further include a radial wall 396 extendingfrom an edge of the vertical wall 394 radially inward toward the axis ofrotation 72. The radial walls 396 of the seats 390, 392 face outwardlyin opposite directions from one another.

The vertical wall 394 of the seats 390, 392 may define one or morecircumferential grooves 398. The Seats 390, 392 may likewise definegrooves 400 extending continuously from vertically along the verticalwall 394 to radially along the radial wall 396. The grooves 398, 404 mayfacilitate the flow of oil around the bearings 370, 372.

Referring to FIG. 18, the fluid lines 132 emanating from the hydraulicrotary union 138 may include feed oil line 410 and a return oil line412. The hub 44 may define a feed port 414 receiving a fitting 416coupling the feed oil line 410 to the feed port 414. Likewise, the hub44 may define a return port 418 receiving a fitting 420 coupling thereturn oil line 412 to the return portion 418. A feed oil passage 422may extend through the hub 44 from the feed port 414 to adjacent theupper bearing seat 390. In the illustrated embodiment, the feed oilpassage 422 is in fluid communication with one or both of thecircumferential groove 398 or the grooves 400 of the upper bearing seat390. In a like manner, a return oil passage 424 extends through the hub44 from the return port 418 to adjacent the lower bearing seat 392. Inthe illustrated embodiment, the return oil passage 424 is in fluidcommunication with one or both of the circumferential groove 398 or thegrooves 400 of the lower bearing seat 392. In the illustratedembodiment, the return oil passage 424 is formed in a ridge 426projecting outwardly from surfaces 428 of the hub 44 having a circularcross section. A corresponding ridge 430 providing a counter balance forthe ridge 426 may be formed opposite the ridge 426 on the hub 44 and thefeed passage 422 may extend through a portion of the ridge 430. Althoughindividual components have been labeled as a “feed” and “return” typecomponents, each of these labels could be reversed for oil flow in theopposite direction.

An upper seal 432 may be interposed between the hub 44 and mast 46 abovethe upper bearing 370 to hinder leakage of oil therefrom. In theillustrated embodiment, an upper cap 434 secures a to the hub 44 inorder to prevent air flow out of the cavity 94. In such embodiments, theupper seal 432 may engage a downwardly depending flange 436 secured toor formed monolithically with the cap 434. The seal 432 also engages themast 46 to create a seal between itself and the flange 436. In someembodiments, the seal 432 may directly engage a corresponding sealingsurface of the hub 44 or some other structure secured to the hub 44.

Referring to FIG. 19, in some embodiments, the return passage 424 may bereplaced with a extends from adjacent the lower bearing seat 292 to apassage 440 extending through the center of the mast 46. The electricallines 130 and fluid lines 132 may also pass through the passage 442. Oil442 passing through the return passage 424 collects within the passage440. In such embodiments, a return oil port 444 coupled to the hydraulicrotary union 138 may conduct oil to the hydraulic rotary union 138,which will conduct the oil away from the passage 440. One or more seals446 engaging the hydraulic rotary union 138 may prevent leakage of theoil past the hydraulic rotary union 138.

Referring to FIG. 20, in some embodiments a lower seal 450 hindersleakage of oil away from the lower bearing 372 between the hub 44 andthe mast 46. In the illustrated embodiment, the upper and lower bearings370, 372 are positioned between the upper seal 432 and the lower seal450. In the illustrated embodiment, the seal 450 secures to a mountingring 452 that is secured to the hub 44. The mounting ring 452 may definea seal seat 454 for receiving the seal 450 and a seal clamp 456selectively secured to the mounting ring 452 by means of a fastener 458,such as a screw, such that the seal 450 is captured between the sealseat 454 and the seal clamp 456. The mounting ring 452 additionallyinclude a circular flange 460 secured to the hub 44 such as by means offasteners 462. The flange 460 may create a sealed interface betweenitself and the hub 44 to hinder oil leakage from between the flange 460and the hub 44. The mounting ring 452 may additionally include one ormore flanges 464. The flanges 464 may extend vertically downward fromthe flange 460. The mast fairing 142 may secure to the flanges 464 bymeans of fasteners 466, such as screws, bolts, rivets, or the like. Inaddition to reducing pressure losses in air flow over the transitionfrom the mast 46 to the hub 44, the mast fairing 142 may additionallyprotect the seal 450 from heated air flow thereover which could degradethe seal 450 and blow past the seal 450 and strip oil from the bearings370, 372.

The seal 450, may include an inner seal 468 and an outer seal 470 thathave identical configurations but mirrored about a horizontal plane. Theseals 468, 470 may have a spacer 472 positioned therebetween. The seals468, 470 and spacer 472 may have the same configuration as the seals340, 342 and spacer 344 of the swashplate 104 as discussed hereinabove.

Referring to FIG. 21, the blade spars 96 may be supported within a sparbore 490 extending through the hub 44 to the cavity 94. An inboardbearing 492 and an outboard bearing 494 may support rotation of theblade spar 96 within the spar bore 490 in order to change the pitchangle 78 of the blade 42 including the blade spar 96. The bearings 492,494 may be embodied as tapered roller bearings. The fluid lines 132emanating from the hydraulic rotary union 138 may include a feed oilline 496 and a return oil line 498 coupled to a feed oil port 500 and areturn oil port 502, respectively. The feed oil port 500 and return oilport 502 penetrate the hub 44 and are in fluid communication with thespar bore 490. An outboard seal 504 and an inboard seal 506 may beinterposed between the blade spar 90 and the spar bore 490 having theinboard bearing 492 and the outboard bearing 494 positioned between theoutboard seal 504 and the inboard seal 506 in order to prevent leakageof oil.

As with other bearings described herein, the bearings 492, 494 mayinclude rolling elements 508 that are preloaded within a certainoperating temperature range. Heating and cooling of the oil passingbetween the feed port 500 and the return port 502 may be used to preventoverheating of the rolling elements 508 due to heated air flow or heatbuildup due to friction and to prevent over reduction or elimination ofthe preload due to thermal shrinkage of the bearings 492, 494, hub 44,or blade spar 96.

The bearings 492, 494 may be protected from heated air flow thereover bythe blade duct fairing 146, which may extend from the blade duct 60 toengage the wall of the cavity 94 of the hub 44. The blade duct fairingmay describe smooth contour from the blade duct 60 to the wall of thecavity 94 such that air flow from the cavity to the blade ductexperiences a smaller pressure drop, such as between five and fiftypercent, preferably between twenty percent and fifty percent, lower thanthe pressure drop that would result if the blade duct fairing 146 wereremoved.

A more complete description of the coupling between the blade spar 96and the hub 44 may be found in U.S. Prov. Pat. App. Ser. No. 61/403,097,filed Sep. 9, 2010 and entitled “FEATHERING-SPINDLE-BEARING LUBRICATIONAND TEMPERATURE CONTROL”.

Referring to FIG. 22, the flow of oil for the lubrication and cooling orheating of the bearings within the rotor 40 may be performed by an oildistribution system 20. The system 20 includes an oil pump 522 poweredby one or both of the engines 48 or by electrical or hydraulic powergenerated by a generator or hydraulic pump powered by the engines 48.The pump 522 may be constantly on or may be turned on and off accordingto conditions of the rotor 40, such as the temperature of one or morebearings thereof. The low pressure port of the pump 522 draws oil from areservoir 524.

The high pressure port of the pump 522 is coupled to a thermalmodulation system 526. The thermal modulation system 526 senses andresponds to the temperature within the rotor system 40. The thermalmodulation system 526 may extract thermal energy from the oil within theoil distribution system 520 in order to lower the temperature ofbearings within the rotor 40 in order to avoid bearing failure due tohigh heat. The thermal modulation system 526 may input thermal energy tothe oil in order to raise the temperature of bearings within the rotor40 in order to avoid thermal shrinkage that will reduce the preload ofrolling elements within the bearings below acceptable levels oreliminate the preload of the rolling elements within the bearingsentirely.

The thermal modulation system 526 may include one or more radiators 528and one or more fans 530 directing air at the radiators 528 in order toextract thermal energy from oil within the oil distribution system. Theradiators 528 are located within the fluid path between the highpressure port and low pressure port of the pump 522.

The thermal modulation system 526 may include one or more heatingelements 532 in thermal contact with oil within the oil distributionsystem 520. The heating elements 532 may be selectively activated toinput heat into oil within the oil distribution system 520. In someembodiments, a bypass valve 534 directs oil to either the radiators 528or heating elements 532 according to the need for heat input to the oilor heat extraction from the oil.

A thermal valve 536 in thermal contact with oil within the oildistribution system 520 may control the fans 530, heating elements 532,and bypass valve 534 according to a temperature of oil within the oildistribution system 520. The thermal valve 536 may be a simplethermostatic switch or may be a digitally programmable sensor andactuator having the capacity to independently control each of the fans530, heating elements 532, and bypass valve 534 in order to modulate thetemperature of oil within the oil distribution system 520.

Oil within the oil distribution system 520 may flow over a swashplatebearing set 538, mast bearing set 540, and spindle bearing set 542. Theswashplate bearing set 538 may include the swashplate bearings 338, 340.Oil flow through the swashplate bearing set 538 may pass through thefeed port 300 a and return port 300 b as described hereinabove. The mastbearing set 540 may include the upper bearing 370 and lower bearing 372.Oil flow through the mast bearing set 540 may include oil flow throughthe feed port 414 and the return port 418. The spindle bearing set 542may include the inboard bearing 492 and the outboard bearing 494. Oilflow through the spindle bearing set 542 may pass through the feed port500 and return port 502 as described hereinabove.

In the illustrated oil distribution system 520, oil flows through theswashplate bearing set 538, mast bearing set 540, and spindle bearingset 542 in parallel through separate fluid paths 544 a, 544 b, 544 c.Temperature controlled valves 546 a, 546 b, 546 c may control oil flowthrough the paths 544 a, 544 b, 544 c, respectively, according to thetemperature of oil exiting the swashplate bearing set 538, mast bearingset 540, and spindle bearing set 542, respectively. Oil flow through thepaths 544 a, 544 b, 544 c may return to the reservoir 524 after exitingthe swashplate bearing set 538, mast bearing set 540, and spindlebearing set 542.

Thermal sensors 548 a, 548 b, 548 c may be in thermal contact with oilflowing through the paths 544 a, 544 b, 544 c downstream from thebearing sets 538, 540, 542, respectively. In some embodiments, anadditional temperature sensor 548 d may sense the temperature of oilflowing from the pump upstream from the radiators 528 and heatingelements 532.

The order of elements along the fluid path between the high pressureport and lower pressure port of the pump 522 may be different that thatillustrated in FIG. 22. Oil distribution systems according toembodiments of the invention may also include only one of a radiator528, or other cooling system, and a heating element 532. Likewise, moreor fewer bearing sets than the swashplate bearing set 538, mast bearingset 540, and spindle bearing set 542 may be lubricated, heated, and/orcooled by oil flowing through the oil distribution system 520.

Referring to FIGS. 23A and 23B, preloading of the bearing setsswashplate bearing set 538, mast bearing set 540, and spindle bearingset 542, may additionally or alternatively be maintained despite lowambient air temperatures, such as during high altitude sustainedlongitudinal flight, by means of heated air forced or drawn through theair channel 92 and the cavity 94.

Referring specifically to FIG. 23A, during sustained longitudinal flightthe rotor hub 44 and blades 42 continue to rotate though unpowered dueto autorotation. The centrifugal force exerted on air within the bladeducts 60 may draw air through the air channel 92 formed by the mastshroud 58 and the mast 46 and into the cavity 94 formed in the hub 44.

In some embodiments, one or more heating elements 550 a, 550 b, 550 care positioned within one or more of the plenum 56 and ducts 54 and areselectively powered to heat air drawn into the mast shroud 58. As notedabove, compressed air from the bypass turbines 62 may be urged throughthe ducts 54 during takeoff, landing, and hover. However, duringsustained longitudinal flight, the engines 48 may operate moreefficiently by directing all bypass air rearwardly from the engines 48rather than through the ducts 54. Accordingly, one or more valves 552 a,552 b may turn off air flow from the bypass turbine 62 to the ducts 54during sustained longitudinal flight of the aircraft 10. However, topermit air flow over the heating elements 550 a, 550 b, 550 c as neededto heat the rotor 40, the valves 550 may be partially opened duringsustained longitudinal flight. In some embodiments, to avoid drawingpower from the engines 48, one or more valves 554 a, 554 b mayselectively permit air flow into the ducts 54 or directly into theplenum 56. The opening and closing of the one or more valves 554 a, 554b may be controlled by temperature feedback from the rotor 40.

For example, a sensor monitoring the temperature of the oil within theoil distribution system 520 may indicate when the oil temperature dropsbelow a certain threshold such that the heating elements 532 are nolonger sufficient to maintain the bearings sets 538, 540, 542 within anoperating temperature range at which the rolling elements thereof arepreloaded, the one or more valves 554 a, 554 b may be partially orcompletely opened and the one or more heating elements 550 a, 550 b, 550c may be activated to warm the rotor 40 to the proper operatingtemperature range at which preloading of the rolling elements within thebearing sets 538, 540, 542 is above a predetermined threshold.

Referring to FIG. 23B, in some embodiments, in lieu of heating elements550 bleed air from a stage of the engines 48 prior to the combustionstage 560 may be conducted by means of one or more channels 562 to oneor both of the ducts 54 or the plenum 56. The channel 562 preferablybypasses any valve 552 a, 552 b controlling bypass air from the bypassturbine 62. Air flow through the channels 562 may be controlled by oneor more valves 564 that may be controlled according to a temperature ofone or more of the bearing sets 538, 540, 542 of the rotor 40. In theillustrated embodiment, the engine 48 includes two compression stages566 a, 566 b and the channel 562 may be positioned to draw air fromeither of the compression stages 566 a, 566 b.

Referring to both FIGS. 23A and 23B, either system may be used to deicethe blades 42 inasmuch as the heated air eventually flows through theblade ducts 60.

Referring to FIG. 24, a flight control system 580 may include flightcontrols 582 providing activating signals, such as electrical,hydraulic, or mechanical inputs, to the swashplate actuators 112, masttilt actuator 174, one or more throttle actuators 584 for controllingthe engines 48 and tip jets 50, and control surface actuators 586 forcontrolling the control surface such as the rudder 22, elevator 30, andailerons 36. The flight controls 582 may receive pilot inputs 588 frompilot controls as known in the art of fixed and rotary aircraft designsuch as rudder control pedals, aileron and elevator control stick,cyclic pitch control stick, throttle control lever, and cyclic pitchcontrol knob. The flight controls 582 may additionally receive inputsfrom an avionic computer 590 enabling autopiloted flying of the aircraft10.

The flight control system 580 may additionally include a thermalmanagement module 592 programmed to maintain the temperature of therotor 40 effective to avoid bearing failure and to maintain bearings andstructures in which they are mounted within an operating temperaturerange in which the bearings will be in a preloaded condition or apreloaded condition above a minimum preload. The thermal managementmodule 592 may receive inputs from the temperature sensors 548 a, 548 b,548 c, 548 d measuring the temperature of oil exiting the swashplatebearing set 538, mast bearing set 540, and spindle bearing set 542,respectively. The thermal management module 592 may be electrically,hydraulically, or mechanically coupled to the valves 552 a, 552 bcontrolling flow of bypass air from the engines 48, the valves 554 a,554 b controlling the air passively drawn into the ducts 54 or plenum56, the valves 564 controlling the flow of bleed air from the engines 48into the ducts 54 or plenum 56, the temperature modulation system 526,including the fans 530 and the heating elements 532, and the thermalvalves 536, 546 a, 546 b, 546 c of the oil distribution system 520. Insuch embodiments, the thermal valves 536, 546 a, 546 b, 546 c may beembodied as electrically, hydraulically, or mechanically actuated valvescontrolled by the thermal management module 592.

The thermal management module 592 may be embodied as a digital or analogcomputer programmed to respond to inputs from some or all of the sensors548 a, 548 b, 548 c, 548 d by activating one or more of the deviceselectrically, hydraulically, or mechanically coupled thereto.Alternatively, the thermal management module may be distributed suchthat each device listed in the preceding paragraph is activated ordeactivated according to a sensed temperature. In particular, thethermal valves 536, 546 a, 546 b, 546 c may respond independently to thetemperature of oil flowing therethrough and open and close according towhether the temperature is within a set operating temperature range,e.g., a temperature range between the temperature at which the rollingelements of the bearing sets 538, 540, 542 will fail and the temperatureat which the preloading of the rolling elements is still present or isabove a proscribed threshold providing the needed rigidity of the rotor40 against destructive harmonics.

FIG. 25 illustrates a method 600 for thermal management of a rotor 40,and, in particular, for reducing probability of bearing failure anddecreasing the amount of slop or play within the swashplate bearing set538, mast bearing set 540, and spindle bearing 542 by maintaining anadequate preload upon the rolling elements thereof. The method 600 maybe performed by the thermal management module 592.

The method 600 may include evaluating at step 602, whether the aircraft10 is taking off, landing, or hovering. The aircraft 10 may be capableof horizontal takeoff along a runway, in which case taking off andlanding for purposes of step 602 may include evaluating whether avertical or short landing or take off is being performed such thatpowered rotation of the rotor 40 by means of the tip jets 50 is neededto achieve the degree of verticality of the landing or take off. Ifhovering, taking off, or landing, is being performed, then at step 604the tip jets are activated at step 604 and compressed air for drivingthe tip jet 50, such as compressed air from the engines 48, is forcedthrough the rotor 40 to the tip jets 50 at step 606. Steps 604 and 606may be performed simultaneously and either step 604 or step 606 may bebegun first. If the aircraft 10 is no longer taking off, landing, orhovering, then the tip jets 50 are deactivated at step 608 and at step610, the compressed air from the bypass turbine 62 is directedrearwardly, in embodiments having engines 48 embodied as jet engines.

Throughout operation of the aircraft, for both sustained longitudinalflight and vertical flight as in a vertical take off or landing orhovering, the method 600 may include executing some or all of steps 612through steps 618.

At step 612 the method 600 includes evaluating whether the temperatureof the rotor 40 is above an upper threshold, such as a temperaturewithin some tolerance of the temperature above which the bearings of theswashplate bearing set 538, mast bearing set 540, and spindle bearing542 will fail or have an unduly shortened useful life. If so, then atstep 614, thermal energy is extracted from the rotor 40. Extractingenergy from the rotor may include activating the fans 530 in order toincrease the rate of heat transfer from the radiators 528.

At step 616, the temperature of the rotor 40 is evaluated with respectto a lower threshold equal to or some tolerance above the temperature atwhich the bearings of the swashplate bearing set 538, mast bearing set540, and spindle bearing 542 are no longer preloaded or have a preloadbelow a minimum preload magnitude. If so, then at step 618, thermalenergy is added to the rotor 40 according to the functionality of theoil and air heating systems described hereinabove. Adding thermal energyto the rotor may include one or more of activating the heating elements532, activating the heating elements 550 a-550 c and opening the valves554 a, 554 b, and opening a valve 564 permitting flow of bleed air overthe rotor 40. The possible methods of adding heat to the rotor 40 may beperformed simultaneously or may be attempted in a specified order suchthat one method is attempted alone, then another method is attemptedsimultaneously if the temperature increase is insufficient, othermethods may then be attempted simultaneously if the temperature increaseis again insufficient.

Referring to FIG. 26, in some embodiments, a method 620 may be used forthermal management of a rotor 40. The method 620 may include evaluatingwhether the average temperature of oil flowing through the rotor 40 hasan average temperature above the operating temperature range at step622. If so, at step 624, heat is extracted from the oil, such as byactivating the fans 530 blowing the radiators 528.

At step 626, the method 620 includes evaluating whether the temperatureof oil circulating through the rotor 40 has a temperature below theoperating temperature range. If so then at step 628, heat is input tothe radiator. Adding thermal energy to the rotor may include one or moreof activating the heating elements 532, activating the heating elements550 a-550 c and opening the valves 554 a, 554 b, and opening a valve 564permitting flow of bleed air over the rotor 40. The possible methods ofadding heat to the rotor 40 may be performed simultaneously or may beattempted in a specified order such that one method is attempted alone,then another method is attempted simultaneously if the temperatureincrease is insufficient, other methods may then be attemptedsimultaneously if the temperature increase is again insufficient.

Steps 622 and 626 may include measuring the temperature of oil enteringor exiting the pump 522 upstream of the swashplate bearing set 538, mastbearing set 540, and spindle bearing set 542. Steps 622 and 626 may beperformed by the thermal valve 536 and steps 624 may include activationof the fans 530 by the thermal valve 536. Opening and closing of one ormore of the valves 554 a, 554 b, 564 and activating of the heatingelements 532, 550 a-550 c may also be controlled according a temperaturedependant signal from the thermal valve 536. Alternatively, or inaddition, activation of the fans 530, opening and closing of the valves554 a, 554 b, 564, and activation of the heating elements 532, 550 a-550c may be controlled by a digital or analog computer, such as the thermalmanagement module 592. In such embodiments, steps 622 and 626 mayinclude evaluating the output of the thermal sensor 548 d. The thermalsensor 548 d preferably measures the temperature of the consolidatedflow of oil from each of the paths 544 a, 544 b, 544 c, such as at apoint between the high pressure port of the pump 522 and the radiators528 and the heating elements 532.

At steps 630, 632, and 634, the temperatures of oil flowing through thepaths 544 a, 544 b, 544 c, respectively, downstream from the swashplatebearing set 538, mast bearing set 540, and spindle bearing set 542 areevaluated to determine whether the temperatures lie within the operatingtemperature range. If the temperature of oil flowing through any of thepaths 544 a, 544 b, 544 c downstream from the swashplate bearing set538, mast bearing set 540, and spindle bearing set 542 is determined tolie outside of the operating temperature range, then at steps 636, 638,and 640 oil flow through whichever of the paths 544 a, 544 b, 544 c hasa temperature outside of the operating temperature range is increased.If the temperature of oil flowing through any of the paths 544 a, 544 b,544 c downstream from the swashplate bearing set 538, mast bearing set540, and spindle bearing set 542 is determined to lie within thepredetermined range, then at steps 642, 644, and 646, oil flow throughwhichever of the paths 544 a, 544 b, 544 c has a temperature outside ofthe operating temperature range is decreased.

Evaluating the temperature of oil flow through the paths 544 a, 544 b,544 c may be performed by the thermal valves 546 a, 546 b, 546 c,respectively. Evaluating the temperature of oil flow through the paths544 a, 544 b, 544 c may additionally or alternatively be performed bythe thermal sensors 548 a, 548 b, 548 c, respectively and the thermalvalves 554 a, 554 b, 554 c may be replaced by valves opened and closedby the thermal management module 592 electrically, hydraulically coupledto the valves in order to increase or decrease the flow of oil throughthe paths 554 a, 544 b, 544 c.

The present invention may be embodied in other specific forms withoutdeparting from its spirit or essential characteristics. The describedembodiments are to be considered in all respects only as illustrative,and not restrictive. The scope of the invention is, therefore, indicatedby the appended claims, rather than by the foregoing description. Allchanges which come within the meaning and range of equivalency of theclaims are to be embraced within their scope.

What is claimed and desired to be secured by United States LettersPatent is:
 1. A method for maintaining rigidity of a rotor system, themethod comprising: providing an aircraft having a rotor system includinga hub defining a cavity, a plurality of blades coupled to the hub andeach defining a duct in fluid communication with the cavity, a pluralityof tip jets secured to the blades in fluid communication with the ducts,a mast having the hub rotationally mounted thereto, a swashplatesurrounding the mast and coupled to the blades and to a plurality ofswashplate actuators, and a shroud surrounding the mast and defining anair channel in fluid communication with the cavity; during at least oneof takeoff and landing, transmitting air having a temperature within afirst temperature range through the rotor system; during sustainedlongitudinal flight, orienting the mast, rotor hub, and blades to beeffective to induce autorotation of the blades; and during sustainedlongitudinal flight, inputting heat to the rotor system effective tomaintain mechanical slack in the relative movement between the mast andplurality of blades within a predetermined tolerance.
 2. The method ofclaim 1, wherein inputting heat further comprises inputting heat to therotor system effective to cause harmonics of the rotor system to lieoutside the range of primary frequencies of cyclic loads induced in therotor system by the blades for flight speeds between 250 and 400 milesper hour and advance ratios higher than 1.0.
 3. The method of claim 1,wherein inputting heat further comprises inputting heat to the rotorsystem effective to cause harmonics of the rotor system to lie outsidethe range of primary frequencies of cyclic loads induced in the rotorsystem by the blades for flight at speeds above 350 miles per hour andadvance ratios higher than 2.0.
 4. The method of claim 1, wherein therotor system further comprises bearings including at least one of: aplurality of spindle bearings interposed between the blades and the hub;a mast bearing interposed between the mast and hub; and a swashplatebearing interposed between a rotating ring and non-rotating ring of theswashplate.
 5. The method of claim 4, wherein inputting heat furthercomprises pumping heated oil over at least one of the bearings.
 6. Themethod of claim 1, wherein inputting heat further comprises: inducingair flow through the air channel, cavity, and ducts by centrifugalforces exerted on air within the ducts due to autorotation of theblades; and heating air flowing through the air channel, cavity, andducts.
 7. The method of claim 1, wherein inputting heat furthercomprises: compressing and heating air within a jet engine; anddirecting the compressed and heated air through the air channel, cavity,and ducts.
 8. The method of claim 7, wherein compressing and heating theair is performed by a bypass turbine.
 9. The method of claim 7, whereincompressing and heating the air is effected by an auxiliary compressorcoupled to the jet engine.
 10. The method of claim 7, wherein directingthe compressed and heated air further comprises directing compressed airfrom the jet engine from a stage located upstream from the combustionstage of the jet engine.
 11. A rotorcraft comprising: an airframe; arotor system comprising a mast mounted to the airframe, a hub rotatablymounted to the mast and defining a cavity, a shroud surrounding the mastand defining an air channel in fluid communication with the cavity, aplurality of blades mounted to the hub, each blade of the plurality ofblades defining a duct in fluid communication with the cavity, and a tipjet mounted to each blade of the plurality of blades to be in fluidcommunication with the duct corresponding thereto; a compressed airsource; a flight control system mounted to the airframe, the flightcontrol system being operably connected to the compressed air source andprogrammed to direct heated compressed air from the compressed airsource through the air channel, cavity, and ducts to the tip jets duringat least one of takeoff and landing, ignite the tip jets only duringsaid directing, and direct the heated compressed air to the rotor systemduring sustained longitudinal flight at a rate effective to maintainmechanical slack in the relative motion between the mast and each bladewithin a predetermined tolerance.
 12. The rotorcraft of claim 11,wherein maintaining the mechanical slack further comprises inputtingheat to the rotor system effective to cause harmonics of the rotorsystem to lie outside the range of primary frequencies of cyclic loadsinduced in the rotor system by the blades for flight speeds between 250and 400 miles per hour and advance ratios of higher than 1.0.
 13. Therotorcraft of claim 11, wherein maintaining the mechanical slack furthercomprises inputting heat to the rotor system effective to causeharmonics of the rotor system to lie outside the range of primaryfrequencies of cyclic loads induced in the rotor system by the bladesfor flight at speeds above 350 miles per hour and advance ratios ofhigher than 2.0.
 14. The rotorcraft of claim 11, wherein the rotorsystem further comprises bearings including at least one of: a pluralityof feathering spindle bearings interposed between the blades and thehub; a mast bearing interposed between the mast and hub; and aswashplate bearing interposed between a rotating ring and non-rotatingring of the swashplate.
 15. The rotorcraft of claim 14, furthercomprising an oil pump and an oil heater in fluid communication with oneor more of the bearings; wherein the flight control system is programmedto activate at least one of the pump and oil heater effective tomaintain slop between the mast and each blade within a predeterminedtolerance.
 16. The rotorcraft of claim 11, further comprising: a plenumin fluid communication with the air channel; a heating elementpositioned within the plenum; and the flight control system, furtherprogrammed to activate the heating element effective to maintain slopbetween the mast and each blade within a predetermined tolerance. 17.The rotorcraft of claim 11, further comprising: a plenum in fluidcommunication with the air channel; a jet engine secured to the airframeand having at least one turbine selectively in fluid communication withthe plenum; the flight control system, further programmed to selectivelydirect heated air from the jet engine to the air channel effective tomaintain mechanical slack in the relative motion between the mast andblades within a predetermined tolerance.
 18. The rotorcraft of claim 17,wherein the at least one turbine comprises a bypass turbine in fluidcommunication with the plenum.
 19. The rotorcraft of claim 17, whereinthe jet engine comprises an auxiliary compressor in fluid communicationwith the plenum.
 20. The rotorcraft of claim 17, wherein the jet enginecomprises a first compressor, a combustion chamber, and a fluid pathbetween the first compressor and combustion chamber; and wherein theplenum is in selective fluid communication with the fluid path.